This invention relates to the field of rocket engine combustion chambers and more particularly relates to combustion chambers having liners supported by jackets and methods of making the same.
The function of a rocket engine main combustion chamber is to contain the combustion process, accelerate the combustion products therefrom to a high velocity and exhaust the combustion products to create thrust. The combustion process occurs at very high temperatures typically at 5,000 to 6,000xc2x0 F. and at high pressures of 1,000 to 4,000 psi. Therefore, it is desirable for the combustion chamber to have a combination of structural strength and an ability to efficiently dissipate heat.
Generally, materials that have thermal conductive properties sufficiently high enough to dissipate the heat of combustion, such as copper, do not possess the structural strength to withstand the pressure of combustion. Therefore, combustion chambers are typically constructed of a combination of materials possessing good thermal conductivity and high structural strength.
Combustion chambers are often constructed of a structurally strong outer shell of steel and a thermally conductive inner liner of copper, or copper alloy. A manifold having coolant channels is defined between the outer shell and the inner liner. The manifold allows liquid coolant to be circulated throughout the combustion chamber for additional heat dissipation. The inner surface of the liner defines a Venturi nozzle in which subsonic combustion gasses are accelerated to supersonic speeds before exiting the combustion chamber. Manufacturing such combustion chambers is typically difficult due to the complex hourglass-shape of the Venturi nozzle and the manifold. Such manufacturing difficulties are further compounded by the use different materials for the outer shell and inner liner.
In one method, the combustion chamber is serially constructed by building up a structural steel jacket around a monolithic inner liner of copper, or copper alloy. The liner is constructed from a roughly cylindrical copper shell which is worked into the Venturi shape including a neck and a pair of flared (bell-shaped) ends to promote combustion, as described above. Wax is applied to the outer surface of the copper shell in a desired configuration for the coolant channels and nickel plating is applied to the wax. After the nickel plating is applied, the liner is completed by melting the wax to leave a network of empty coolant channels defined between the nickel plating and the cooper shell. Several individual steel plates are then welded together in a conforming fit around the liner so as to form an outer structural jacket.
Advantageously, the serially constructed liner has high thermal conductive properties and the jacket provides sufficient strength to withstand the pressures generated during the combustion process. The coolant channels allow liquid coolant to be dispersed through the liner during combustion to remove heat from the combustion chamber. However, the number of welds and other labor-intensive activities required to form the structural jacket from the steel plates increases the cost of constructing the combustion chamber. Further, the serial process of building the combustion chamber is prone to failure because a single mistake in one of the steps will result in scrapping of the entire piece. This is especially costly as the combustion chamber nears completion and a significant amount of labor and materials have been invested in the production process.
Another method of manufacturing a combustion chamber is disclosed in U.S. Pat. No. 5,701,670 to Fisher et al (xe2x80x9cFisherxe2x80x9d). Fisher discloses a method of making a rocket engine combustion chamber that uses three basic components including a structural jacket, a monolithic coolant liner, and a plurality of throat support sections. The liner is a copper shell machined into the Venturi shape and includes coolant channels preformed on its outer surface. The throat support sections are fabricated and assembled around the indentation created by the tapered neck of the combustion chamber liner. The structural jacket is a heavy cast metal cylinder defining a cylindrical opening.
The throat support/liner subassembly is installed into the cylindrical opening of the structural jacket using a shrink-fit process. The shrink-fit process involves chilling the throat support/liner subassembly to a cryogenic temperature and heating the structural jacket. Heating of the structural jacket and cooling of the support/liner subassembly increases the clearance within the jacket. This increase in clearance allows insertion of the throat support/chamber liner subassembly into the structural jacket. The entire combustion chamber assembly is then subjected to a hot-isostatic pressure (HIP) bonding process. The HIP bonding process includes heating and pressurizing the entire assembly until the copper shell softens and adheres to the inner surfaces of the adjacent throat supports or the structural jacket.
The combustion chamber disclosed in Fisher has the advantage of lower cost due to a less labor-intensive construction and non-serial construction that is more forgiving of mistakes. However, in order to provide sufficient clearance to accept the throat support/liner subassembly, the structural jacket must be roughly cylindrical. The cylindrical jacket is much heavier than the jacket used in serial construction which can be more closely fit to the Venturi shape of the liner. Further, the shape difference between the cylindrical passage of the jacket and the liner requires the use of the thick and heavy throat supports. These throat supports must be heavy and strong to resist the increased pressures that occur during the HIP bonding process. The increased weight of such a combustion chamber has the effect of reducing the payload of the rocket.
Therefore, it would be advantageous to provide a relatively easy, inexpensive method of making a combustion chamber that still has sufficient structural integrity to withstand the heat and pressure of combustion. It would also be advantageous if the combustion chamber were relatively light so as to minimize its impact on the payload capacity of the rocket.
The present invention addresses the above needs and achieves other advantages by providing a combustion apparatus for containing and directing the combustion of a propellant that includes a structural jacket defining a contoured passage having a forward liner positioned within a first end of the passage and an aft liner positioned within a second end of a passage. Advantageously, a liner joint between the forward and aft liners is positioned downstream of a throat of the forward liner where the combustion gasses have a considerably lower heat flux. In addition, the contoured passage of the structural jacket eliminates the need for throat supports which reduces the weight of the combustion chamber and the complexity of assembly of the combustion chamber.
In one embodiment, the present invention includes a combustion assembly for containing and directing combustion of a propellant. The combustion assembly includes a structural jacket defining a passage including a first end, a second end and a neck. The neck is positioned between and separates the first and second ends of the passage. A forward liner is configured to fit with the first end of the passage and includes a throat and a downstream portion. The throat is upstream of the downstream portion and is configured to fit within the neck of the passage. Configured to fit within the second end of the passage is an aft liner. The aft liner includes an upstream portion configured to abut the downstream portion of the forward liner so as to form a liner joint. When fit into the structural jacket, the forward and aft liners cooperate to form a longitudinal combustion chamber so that combustion of the propellant is contained and directed through the combustion chamber from the forward liner to the aft liner.
In another aspect, each of the upstream and downstream portions includes a joining flange, wherein the joining flanges are configured to abut each other in an overlapping relationship. Each of the upstream and downstream portions may also include a beveled portion partially formed by the flanges and configured to abut each other in an overlapping relationship so as to form the liner joint. The flanges allow the liners to be secured to each other during a bonding process that seals the liners to the structural jacket. The excess portions of the flanges may be removed after bonding for a smooth, contiguous liner surface. One of the flanges may also include a HIP assist groove that decreases the contact area between the flanges which increases the pressure at the interface between the flanges during the bonding process.
In yet another aspect, the neck of the passage may be over-bored to allow passage of the joining flange of the forward liner. Optionally, the assembly can also include a filler ring configured to fit around the throat of the forward liner. The filler ring forms a close fit between the throat and the neck after insertion of the forward liner into the jacket. Preferably, the filler ring is constructed of a brazeable material allowing it to be brazed into place between the throat and the neck.
In still another aspect, one of the joining flanges includes a pin and the other one of the joining flanges defines an index hole configured to receive the pin. The index hole has a width equal to a diameter of the pin and a depth at least as great as a length of the pin. The depth of the index hole allows the flanges to come into full contact. Use of the pin and index hole ensures that coolant channels on the surface of the liners are properly aligned.
The combustion chamber apparatus of the present invention has several advantages. The use of a two or more liner sections in lieu of a single piece liner allows larger combustion chambers to be constructed from standard sized liner shells, billets or casings. The need for throat supports is eliminated by the contoured shape of the structural jacket. Eliminating the throat supports has several advantages, including reduction in the overall weight of the combustion chamber, elimination of the occurrence of coolant leakage between the throat supports and simplification of the assembly process. Further, the wall structure of the jacket may be thinner, and hence lighter, due to the conforming shape of the jacket passage and the ability to directly apply pressure during the HIP bonding process to the inner wall of the jacket through the hole defined in the outer wall if a throat support cylinder is used. Optionally, some chambers can be constructed without an outer wall, or the outer wall may be removed entirely. A lighter wall structure, combined with elimination of the heavy throat supports, results in an overall decrease in the weight of the combustion chamber and a concomitant increase in the payload of the rocket.